Aircraft navigation-aid apparatus



Dec. 17, 1957 c. A RlcHARDsoN AIRCRAFT NAVIGATION-AID APPARATUS FiledDeo. 8, 1953 ICS AIRCRAFT N AVIGATION-AID APPARATUS Application December8, 1953, Serial No. 396,837

Claims priority, application Great Britain December 19, 1952 5 Claims.(Cl. 343-108) This invention relates to aircraft flight controlapparatus for use in facilitating the steering of an aircraft and, moreparticularly, for use in facilitating the safe landing of an aircraftalong an inclined landing path or surface, such as a glide-path definedby radio beams.

The invention relates more specifically to such apparatus of the kind inwhich pitch guidance apparatus for use in controlling the flight of theaircraft in pitch receives as a control signal the combination ofsignals in appropriate senses including one, derived from radioapparatus, that is a measure of the vertical displacement of theaircraft from the landing path, and another, derived from apitch-change-sensitive device, that is a measure of the angulardeviation in pitch of the aircraft from a preset attitude in pitch.

The pitch guidance apparatus may be, or may form part of, an indicatinginstrument having an indicator, such as a horizontal pointer, movablerelative to a fixed reference index under the control of the ightsignal, from which the pilot may guide his aircraft in pitch to proceedalong the landing path, or it may be the elevator servo motor of anautomatic control system for the aircraft which is energised by thecontrol signal to control the aircraft automatically in pitch to proceedalong the landing path.

Aircraft flight control apparatus in which the pitch guidance apparatusforms part of an indicating instrument is disclosed in Patent No.2,654,086 to Cecil C. Pine and Charles L. Sharp for Safety Device forInstrument Approach Systems, dated September 29, 1953. With the type ofight control apparatus disclosed in that specification the pilot hasonly to control his aircraft so as to maintain the indicator at a zeroposition with respect to the fixed reference index in order to ensurethat, at least in the absence of a vertical component of wind, theaircraft will proceed down the landing path, or, if vertically displacedfrom it, will approach it asymptotically in a vertical plane. Suchapparatus has now become Well-known and has proved extremelyadvantageous to pilots in cross-country flight, in flying on adirectional radio beam, and in ying down a landing beam. Broadly,speaking, such apparatus comprises a crossed pointer meter in which twopointers-a horizontally movable vertical pointer and a verticallymovable horizontal pointer-move over a common dial from zero referencepositions crossing in the centre of the dial. These pointers arecontrolled by various combinations of signals according as the aircraftis to be controlled to ilyv in one or another of several ight pathsdened in different ways. In each of the alternative combinations ofsignals one signal measures the linear or angular displacement of theaircraft from the prescribed path or direction, and one or more of theother signals measures one or more time derivatives of that displacment.Whichever combination of signals is used to control the pointers, theaforesaid method of controlling the aircraft in tlight is used, that is,the aircraft is controlled to y aims in such a way that the two pointersare maintained at their zero reference positions.

If it is desired to control the aircraft to proceed down a landing pathdened by localiser and glide-path beams, the vertically movablehorizontal pointer is controlled from a combination of signals includingone, derived from a glide-path receiver, that is a measure of thevertical displacement of the aircraft from the glidepath, and another,derived from a pitch-change-sensitive device, that is a measure of theangular displacement of the aircraft in pitch from a preset attitude inpitch. The latter signal may be considered as representing the firsttime derivative of the vertical displacement from the glide path. Thehorizontally movable vertical pointer is controlled by a combination ofsignals including one derived from a localiser receiver that is ameasure of the lateral displacement of the aircraft from the localiserpath, another derived from a direction-giving instrument that is ameasure displacement of the aircraft in azimuth from the direction ofthe localiser path, and a third derived from an attitude-indicatinginstrument that is a measure of the bank angle of the aircraft. The tWolast-mentioned signals may be considered as representing the first andsecond time derivatives of the lateral displacement of the aircraft fromthe localiser path.

It will be appreciated that flight control apparatus of the kinddescribed to which the present invention is applicable comprises variousauxiliary apparatus such as amplifiers, signal modifiers, .combiners andlimiters, together with a control panel which includes a selector switchby ywhich the pilot is enabled to select a combination of signalsappropriate for the type of iiight he Wishes to carry out. His controlpanel may also include various other control knobs such as a pitch-trimknob and an altitude control knob.

As has been stated, the pitch-guidance apparatus may form part of anautomatic control system for aircraft. An automatic control system ofthis kind is also disclosed in Patent No. 2,611,128 to Cecil C. Pine andCharles L. Sharp for Safety Device for Automatic Approach Systems, datedSeptember 16, 1952.

It may be that, during a landing procedure when an aircraft is beingcontrolled either manually from an indicating instrument orautomatically through an automatic control system, the pilot mayconsider, for some reason or another, that his approach to the landingbase will not be a satisfactory one, and that therefore he will not wishto land but to iy around the landing base. In such an event it will beappreciated that it is extremely desirable that the pilot should beenabled to modify the controls to the indicating instrument or to theautomatic pilot so that he may still be able to make use of them tocause the aircraft to be guided either manually or automatically awayfrom the landing base. ln fact it is eX- tremely desirable that themodifying controls should be such that if the aircraft is controlled bythe indicator or the automatic pilot in the usual manner, the aircraftwill proceed to climb away from the landing base.

A proposal has been made in specification No. 691,017 to provide, inapparatus of the kind referred to, emergency means which are operativeto render the pitchguidance apparatus responsive only to the signal fromthe pitch-change-sensitive device, and which is also operative to rendereffective a normally ineffective biasing device that biases the signalfrom the pitch-change-sensitive device in such a manner that thepitch-guidance apparatus receives an operative signal calling for aclimb of the aircraft at a predetermined angle.

It is an object of the present invention to provide an 0 improvement inaircraft flight control apparatus of the kind described in the aforesaidPatents 2,654,086 and 2,611,128.

The desired angle at which an aircraft should climb away from a landingbase whenA the emergency means has been rendered effective variesaccording to a number of factors, an important one of which is theposition of the wing-aps of the aircraft.

According to the present invention there is provided flight controlapparatus of the kind claimed in any one of claims l to 13 ofspecification No. 691,017 including means to ensure that the signal fromthe pitch-changesensitive device is biased in such a manner that theoperative signal calling for a climb of the aircraft at a predeterminedangle is dependent on the flap angle.

On most types of aircraft the ap angle is shown to the pilot by a smallindicator on or adjacent to his instrument panel which is operated by acontrol signal provided by a device for measuring the flap angle.According to a particular embodiment of the present invention thissignal that operates the ap angle indicator together with the normalpitch-deviation signal and a biasing signal from the biasing device aresupplied in the appropriate senses to control the horizontalvertically-movable pointer when the emergency means is renderedeffective.

One form of the invention will now be described by way of example withreference to the accompanying drawings wherein:

Figure l is a diagrammatic arrangement of a flight control systemembodying the invention, and

Figure 2 shows how the switch which transfers the pitch angle controlmay be associated conveniently with the engine throttle lever.

Figure 3 shows the selector switch usually used in connection with theknown zero reading guidance system to which this invention relates.

In the embodiment shown, during radio guided approaches, the output ofthe radio glide path receiver 1 after passing through the modulator 2vis led to a mixer amplifier and rectifier 3 where this signal iscombined with the stabilising pitch signal from the gyro-vertical 4. Thepitch signal is represented as generated by a synchro 5 on the pitchaxis of gyro 4, which produces a signal proportional in amount to thedeparture of the craft from a level or trim position and reversible inphase when such departure is up or down. The combined A. C. signal isamplified and rectified and the output led to the zero reading indicator6 thereby causing movement of the horizontal pointer 7 up or down fromits zero or central position, according to whether the aircraft needstrimming up or down to keep it on the glide path.

As explained in the patent to Spencer Kellogg 2d, No. 2,613,350 forFlight Indicating System for Dirigible Craft, dated October 7, 1952, itis characteristic of this type of meter that it indicates zero not onlywhen the aircraft is flying on the glide path (or level), but also whenthe aircraft has departed from the glide path as soon as its trim hasbeen changed proportionally to the amount of departure to bring theaircraft smoothly back into the glide path.

Similarly, the output of the localiser receiver 8 is led through amodulator 9 and to a mixer amplifier and rectier 10 where the signal iscombined with two other signals, one a course signal from the synchro 11operated from the compass 12 and the other as signal from the synchro 13operated from the bank axis of the gyro-vertical 4. The combined outputof the three signals is fed to the indicator 6 to cause lateral movementof the pointer 14 to the right or left from its normal or zero position.As with the case of the pointer 7, the pointer 14 will read zero notonly when the craft is on the radio course, but also is off course whenthe craft is banked at a proportional angle proportionate to coursedeparture to cause the craft to return smoothly to its radio course.

An emergency switch or push button 15 is provided which the aviator maypress or throw, in case he desires not to land when approaching thetouchdown point, but

decides to go around again. The pressing of pushlrfuttonA 15, it will beseen, will sever the output of the glide path modulator Z from the mixerrectifier 3, thus leaving thesignal from the attitude gyro in solecontrol to erase the downward glide signal. The gyro signal is alsomodied to call for a climb, by throwing in a biasing means such as asignal from a synchro 16 which is given a predetermined displacementfrom knob 17 so as to set in a predetermined pitch signal into themixer, thus causing the indicator to immediately call for a climb at asafe angle.V As the engine speed should be increased simultaneously, wehave shown the button as conveniently placed on the engine throttle 18(Fig. 2), so that the two operations may be accomplished as one.

The same button 15 also preferably either breaks the circuit from thelocalizer modulator or modifies the signal therefrom to prevent erraticoperation of the meter as the craft closely approaches the localisertransmitter. Pressing the button, therefore, also first opens thebridging contact 35 in the circuit between modulator 9 and the receiver10 and closes a circuit between contacts B and B thereby placing asignal attenuator in the shape of a variable resistor 19 in the circuit.The magnitude of the resistance is adjusted to such a value that theaircraft, when controlled in accordance with the indicator, neverexceeds an angle of bank of 10, thus avoiding excessive banks while thecraft is near the ground. The button is also shown as having a thirdcontactor 20 which throws into the circuit a device 25 for reducing theamplitude of the signals from the localiser receiver as the transmitteris being approached. Such an operation is termed'in the art coursesoftening and is described in the specification of U. S. Patent No.2,439,044 to Thomas M. Ferrill, Jr., for Coarse Softening System, datedApril 6, 1948. A- third position of the switch is shown, in which thelocaliser receiver is completely disconnected. Under such condition, thecraft will still be kept on course by following the zero readingindicator, and if course changes are desired the synchro 11 could beadjusted from the knob 21.

As soon as the craft is safely away from the landing fieldV the pilot,by operating his selector switch 30, may

resume the flight system desired by restoring switch 1S to its normalposition and by moving the course selector handle 31 to 'the flightinstrument position for regular cross country flight or to the ODRposition for radio guided cross country flight or resume the approachposition, in which case, by following the indicator 6, the pilot willagain causethe craft to approach and reach a landing position over thelanding field runway, all as more fully described in the aforesaidPatent No. 2,613,350

When the pilot decides-to forego landing andclimb, he adjusts his enginespeed and presses the button 15 to cut out the guide path control signaland the signal from the localiser receiver 8 and inserts into the mixerrectifier 3 a predetermined signal from the synchro 5 together with asignal derived from the angular position of the flaps by way of theflap-angle indicator 36'.

If, as is usually the case just before landing, the aps are in anoperative position, the climbing angle of the aircraft is greater thanit would be with the flaps fully retracted and the attitude of the bodyofthe wing unchanged.

It follows, therefore, that in order toy maintain unchanged a desiredangle of climb of the aircraft the'signal derived from the flap anglemust be arranged to increase with decrease of that angle, reaching aymaximum when the Hap is fully retracted, and must at all times beV of aValue which, together with the biassing signal from the synchro, keepsthe total signal to the mixer constant in magnitude, and of a valueconsonant with a predetermined `angle of climb-of the aircraft.

Byincorporatingthefeature of the present invention in' aircraft flightcontrol apparatus of the'kindreferredtoit is ensured that `the initialclimb-away 'angle ofthe aircraft is suitableto the flapV positionexisting at the timeof'rendiw-V ing the emergency means effective and,as the aps are retracted during the climb-away procedure, the climb-awayangle is modified. In this Way an additional safety precaution isintroduced into the apparatus.

I claim:

1. A radio approach system for aircraft for facilitating the landing ofthe aircraft in a landing field comprising pitch-guidance apparatus, alanding beam receiver, a pitch-deviation measuring instrument, means forcombining a signal derived from said receiver representing the verticaldisplacement of the aircraft from a radio-defined landing path with asignal derived from said pitchdeviation measuring instrumentrepresenting the angular deviation of the aircraft from a predeterminedattitude in pitch and for supplying said combined signals to operate thepitch-guidance apparatus, characterised in that there is providedemergency means which is operative to render the pitch-guidanceapparatus responsive only to the signal from the pitch-deviationmeasuring instrument and which is also operative to render effective anormally ineffective biasing device that biases the signal from thepitchdeviation measuring instrument in such a manner that thepitch-guidance apparatus receives an operative signal calling for aclimb of the aircraft at a predetermined angle, and means for modifyingsaid signal when the ap angle on the craft is other than normal.

2. In a guidance and navigational system for aircraft including a radioIglide path guidance receiver, a pitch measuring device, a null readingindicator actuated jointly by said receiver and device, for facilitatingthe landing of the aircraft as claimed in claim 1, in which biasingmeans 6 are provided for lessening the climbing attitude of the craftwhen the flaps are down.

3. A system as claimed in claim 2 in which the biasing device produces asignal which is combined with the signal from the pitch-deviationmeasuring instrument and supplied to the pitch-guidance apparatus as afly-up control signal.

4. In a guidance and navigational system for aircraft including a radioglide path guidance receiver, a pitch measuring device, a null readingindicator actuated jointly by said receiver and device, for facilitatingthe landing of the aircraft, a safety device for disconnecting at leastthe glide path radio receiver at will, including means for causing saidpitch measuring device to set the indicator to call for a set climbingattitude, and means for automatically varying the indicated attitudewhen the angle of the flaps on the craft is other than normal.

5. In a blind instrument landing system for aircraft, a null or zeroreader pitch attitude indicator, an attitude gyro producing a pitchsignal, a radio landing beam receiver producing a signal, said indicatorbeing normally actuated by both of said signals, means for biasing saidpitch signal and disconnecting said receiver signal in an emergency, andmeans for altering said biased signal when the flap angle on the craftis other than normal.

References Cited in the le of this patent UNITED STATES PATENTS2,611,128 Pine et al. Sept. 16, 1952 2,654,086 Pine et al Sept. 29, 19532,683,004 Alderson et al. July 6, 1954

